Metallurgical product and structure member for aircraft made of Al-Zn-Cu-Mg alloy

ABSTRACT

The present invention relates to a work-hardened product, and particularly a rolled, extruded and/or forged product made of an alloy with composition (% by weight): Zn 6.7-7.3% Cu 1.9-2.5% Mg 1.0-2.0% Zr 0.04-0.15% Fe ≦0.15 Si ≦0.15 other elements ≦0.05 each and ≦0.15 total, remainder aluminium and wherein Mg/Cu&lt;1. The product is preferable treated by solution heat treatment, quenching, cold working and artificial aging. Cold working may be achieved by controlled tension and/or cold transformation, for example rolling or drawing. The product may be used, for example, as an aircraft structural member.

CROSS REFERENCE TO RELATED APPLICATIONS

The present application claims priority to U.S. Ser. No. 60/529,594filed Dec. 16, 2003, the content of which is incorporated herein byreference in its entirety.

BACKGROUND OF THE INVENTION

1. Field of the Invention

The present invention relates generally to rolled, extruded and/orforged products made of Al—Zn—Cu—Mg alloy treated by solution heattreatment, quenching, cold working and artificial aging, andparticularly structural members made from such products and designed foruse such as in aircraft construction.

2. Description of Related Art

It is generally known that when manufacturing semi-finished products andstructural members for aeronautical construction, certain variousrequired properties are difficult to optimize at the same timeindependently of each other. When the chemical composition of the alloyor the parameters of production processes are modified, several criticalmechanical properties can tend to vary in opposite directions. This issometimes the case with respect to properties collectively known in theart as “static mechanical strength” (particularly the ultimate tensilestrength R_(m) (or UTS) and the tensile yield strength R_(p0.2)(orTYS)). Other properties also at issue are commonly referred to under theterm “damage tolerance” properties (particularly toughness andresistance to crack propagation). Some frequently employed propertiessuch as fatigue strength, corrosion resistance, formability andelongation at failure are related to these mechanical properties (or“characteristics”) in a complicated and frequently unpredictable manner.Therefore, optimization of all properties of a material for aeronauticalconstruction very frequently means making a compromise between severalkey parameters.

Typically, 7xxx type alloys are used for wing structural members (exceptfor undersurface wing members).

U.S. Pat. No. 5,865,911 (Aluminum Company of America) discloses anAl—Zn—Cu—Mg type alloy with composition:

-   -   Zn 5.9-6.7, Mg 1.6-1.86, Cu 1.8-2.4, Zr 0.08-0.15        for making structural members for aircraft. These structural        members are optimized to have high mechanical strength,        toughness and fatigue strength.

Patent application WO 02/052053 describes three Al—Zn—Cu—Mg type alloyswith composition (a) Zn 7.3+Cu 1.6; (b) Zn 6.7+Cu 1.9; (c) Zn 7.4 Cu1.9; each of these three alloys also containing Mg 1.5+Zr 0.11. This WOpublication also describes appropriate thermomechanical treatments formaking structural members for aircraft.

Furthermore, a 7040 alloy with the following standardized chemicalcomposition is known: Zn 5.7-6.7 Mg 1.7-2.4 Cu 1.5-2.3 Zr 0.05-0.12 Si ≦0.10 Fe ≦ 0.13 Ti ≦ 0.06 Mn ≦ 0.04other elements ≦0.05 each and ≦0.15 total.

A 7475 alloy with the following standardized chemical composition isalso known: Zn 5.2-6.2 Mg 1.9-2.6 Cu 1.2-2.9 Cr 0.18-0.25 Si ≦ 0.10 Fe ≦0.12 Ti ≦ 0.06 Mn ≦ 0.06other elements ≦0.05 each and ≦0.15 total.

Alloys in the 2xxx series are routinely used, for example the 2324alloy, for some structural members of civil aircraft wings such as underwing members.

Alloys conventionally used for fuselage structural members typicallybelong to the 2xxx series, for example the 2024 alloy.

SUMMARY OF THE INVENTION

A purpose of the present invention was to obtain aircraft structuralmembers, and particularly fuselage members made of Al—Zn—Cu—Mg alloy,with a higher mechanical strength than is possible in prior alloys, withcomparable damage tolerance and sufficient formability.

Another purpose was to obtain aircraft structural members, andparticularly members for the under surface wings of aircraft, or formachining integral structures made of Al—Zn—Cu—Mg alloy, with a bettercompromise between mechanical strength, toughness and fatigue strengthproperties, than is possible to achieve with prior materials.

In accordance with these and other objects there is provided awork-hardened product, (preferably a rolled, extruded and/or forgedproduct) of an alloy comprising (% by weight):

-   Zn 6.7-7.3% Cu 1.9-2.5% Mg 1.0-2.0%-   Zr 0.04-0.15% Fe ≦0.15 Si ≦0.15    other elements ≦0.05 each and ≦0.15 total, the remainder being    aluminium, wherein Mg/Cu<1. The product is preferably treated by    solution heat treatment, quenching, cold working and artificial    aging. Cold working may be achieved by controlled stretching and/or    cold transformation, for example by rolling or drawing.

In further accordance with the present invention there is provided astructural member suitable for aeronautical construction, andparticularly for an aircraft fuselage, or for members of the undersurface of an aircraft wing, or an integral structural member for anaircraft, made from such a work hardened product, and particularly fromsuch a rolled or extruded product.

Additional objects, features and advantages of the invention will be setforth in the description which follows, and in part, will be obviousfrom the description, or may be learned by practice of the invention.The objects, features and advantages of the invention may be realizedand obtained by means of the instrumentalities and combinationparticularly pointed out in the appended claims.

DETAILED DESCRIPTION OF A PREFERRED EMBODIMENT

Unless mentioned otherwise, all information about the chemicalcomposition of alloys is expressed in percent by mass. Consequently, ina mathematical expression “0.4 Zn” means 0.4 times the zinc contentexpressed in percent by mass; the same applies by analogy to otherchemical elements. Alloys are named in accordance with the rules of TheAluminium Association, known to those skilled in the art. Themetallurgical tempers are defined in European standard EN 515. Thechemical composition of normalised aluminium alloys is defined forexample in standard EN 573-3. Unless mentioned otherwise, staticmechanical characteristics, in other words the ultimate tensile strengthRm, the tensile yield strength R_(p02) and the elongation at fracture A,are determined by a tensile test according to standard EN 10002-1, thelocation at which the pieces are taken and their direction being definedin standard EN 485-1. The fatigue strength is determined by a testaccording to ASTM E 466, and the fatigue crack propagation rate (usingthe da/dn test) according to ASTM E 647. The R curve is determinedaccording to ASTM standard 561. The critical strength intensity factorK_(C), in other words the intensity factor that makes the crackunstable, is calculated starting from the R curve. The stress intensityfactor K_(CO) is also calculated by assigning the initial crack lengthto the critical load, at the beginning of the monotonous load. These twovalues are calculated for a test piece of the required shape. K_(app)denotes the K_(CO) factor corresponding to the test piece that was usedto make the R curve test. Unless otherwise mentioned, the crack size atthe end of the fatigue precracking stage is W/4 for test pieces of theM(T) type, and W/2 for test pieces of the CT type, wherein W is thewidth of the test piece as defined in standard ASTM E561.

The term “extruded product” includes so-called “drawn” products, inother words, products produced by extrusion followed by drawing.

Unless mentioned otherwise, the definitions in European standard EN12258-1 are applicable.

The term “structural member” in this specification refers to amechanical part used in mechanical construction, for which failure couldendanger the safety of the said construction and/or its users, orothers. For an aircraft, these structure members include particularlymembers making up the fuselage (such as the fuselage skin, fuselagestiffeners or stringers, bulkheads, fuselage circumferential frames,wings (such as wing skin), stringers or stiffeners, ribs and spars andthe tail fin composed particularly of horizontal and verticalstabilisers, and floor beams, seat tracks and doors.

For the purposes of this description, “integral structure” means thestructure of part of an aircraft that was designed to achieve materialcontinuity over the largest possible size in order to reduce the numberof mechanical assembly points. An integral structure may be made eitherby in-depth machining, or by the use of shaped parts for exampleobtained by extrusion, forging or casting, or by welding of structuralmembers made of weldable alloys. Thus, the result is large single-piecestructure members, with no mechanical assembly or with a small number ofmechanical assembly points compared with an assembled structure in whichthe thin or thick plates depending on the destination of the structuremember (for example fuselage member or wing member) are fixed, usuallyby riveting, onto stiffeners and/or frames (that can be made bymachining from extruded or rolled products).

The present invention can advantageously be applied to an aluminiumalloy containing from about 6.7% to about 7.3% of zinc. The zinc contentshould preferably be high enough to achieve good mechanical properties,but if it too high, the sensitivity of the alloy to quenching mayincrease, which in particular introduces a risk for thick products ofdegrading the compromise between target properties. In one advantageousembodiment of the instant invention, the product is a plate thinner thanabout 20 mm. In another advantageous embodiment of the invention, theproduct is a thick plate, thicker than about 20 mm.

The chemical composition of the Al—Zn—Cu—Mg alloy was chosen such thatthe Mg/Cu ratio of the alloy according to the invention is preferablybelow about 1. Preferably, this ratio is kept at a value less than 0.9.A value less than 0.85, or even about 0.8 is preferred.

An advantageous compromise was found when the copper content was kept atfrom about 1.9 to about 2.5%, and preferably from about 2.0 to about2.3%, while the content of magnesium is fixed at from about 1 to about2% and preferably from about 1.5 to about 1.8%.

The inventors have observed that a zirconium content from about 0.07 toabout 0.13% made it possible to achieve a better compromise betweenR_(P0.2), toughness (at ambient temperature or when cold) and fatiguestrength (particularly the propagation rate of fatigue cracks), for thiscomposition of major elements Al—Zn—Cu—Mg. If the content of Zr exceedsabout 0.12%, there may be a significant risk of primary Al₃Zr typephases being formed, unless cooling is fast enough; in the case ofsemi-continuous casting, such a sufficient rate can be achievedparticularly when billets are being cast.

The Zr content for rolled products is preferably less than about 0.12%,and advantageous results have been obtained with a content of from about0.07 to about 0.09%. A zirconium content of up to about 0.13% can besuitable for billets in some embodiments.

In all cases, silicon and iron contents should preferably each be keptbelow about 0.15% and particularly preferably below about 0.10% to havegood toughness. In one particularly preferred embodiment of the presentinvention, the iron content preferably does not exceed about 0.07%, andthe silicon content preferably does not exceed about 0.06%.

An alloy according to the instant invention can be cast according to oneof the techniques known to those skilled in the art to obtain unwroughtproducts such as an extrusion billet, or a rolling ingot. This unwroughtproduct, possibly after scalping, is then homogenized, typically for aduration of 15 to 16 hours at a temperature of preferably from about 470to about 485° C.

The unwrought product is then transformed hot into extruded products(particularly bars, tubes or sections), hot rolled plates or forgedparts. In one preferred embodiment of the invention, the inventors haveobserved that surprisingly, thick products according to the inventioncould be hot rolled at a temperature of about 350° C., which is muchlower than the temperature usually used for this type of product (whichis about 415 to 440° C.) without affecting the required compromisebetween properties for thick products used in aircraft structures.

Hot transformation may possibly be followed by cold transformation. Forexample, extruded and drawn tubes can be manufactured. In the case ofrolled products, one or several cold rolling passes may also beperformed. This may be necessary if the desired final thickness is belowabout 3 mm.

Products obtained are then preferably solution heat treated. Thissolution heat treatment may be made in any appropriate furnace such asan air furnace (horizontal or vertical) or a salt bath furnace. In onepreferred embodiment of the invention leading to thick products (>10mm), this solution heat treatment is carried out at a temperature fromabout 470 to 480° C. and preferably from about 475 to about 480° C.preferably for at least 4 hours. In another preferred embodiment,leading to thin products (<10 mm), the solution heat treatment iscarried out at preferably from about 470° C. to about 475° C., and theduration of the solution heat treatment, for which the optimum valuedepends on the product thickness, is typically at least about one hour.

The products are then quenched, preferably in a liquid medium such aswater, the liquid preferably being at a temperature of not more thanabout 40° C.

The products are then usually subjected to controlled stretching with apermanent set of the order of preferably from about 1 to 5%, andparticularly preferably from about 1.5 to 3%.

Finally, the products are subjected to an artificial aging treatmentthat has a large influence on the final properties of the product.Depending on the required compromise, a two-step artificial aging or asingle step artificial aging may be preferred.

Products according to the invention result in new products withparticularly attractive characteristics for aeronautical construction.These products may be in any desired form such as in the form of sheetor plates, particularly fuselage sheet, thick plates for undersurfacestructural members or for integral structures, or in the form ofextruded sections, or in the form of forged parts. Rolled productsaccording to the present invention may be thick or thin and have atensile yield strength R_(p0.2(L)) equal to at least about 500 MPa andpreferably at least about 520 MPa and even at least about 530 MPa, for aK_(app(L-T)) measured according to ASTM 561 on a C(T) type test piecewith W=127 mm and thickness B=5.5 mm equal to at least 100 Mpa{squareroot}m, or possibly even more than 110 Mpa{square root}m.

In one advantageous embodiment of the present invention, sheet and platewith a thickness from about 1 to about 10 mm has a K_(C(T-L)) value,measured on a test piece with W=760 mm, of at least about 130 Mpa{squareroot}m, and/or a K_(C(L-T)) value, measured on a test piece with W=760mm, of at least about 160 Mpa{square root}m.

In another advantageous embodiment of the invention, the thickness ofthe sheet or plates obtained is more than or equal to about 20 mm, andthe sheet or plate has a yield strength R_(p0.2)(L) of at least about520 MPa, a K_(app(L-T)) measured according to ASTM E 561 on a C(T) typetest piece with W=406 mm and thickness B=6.35 mm equal to at least about130 Mpa{square root}m, and K_(C(L-T)) measured on a C(T) type test piecewith W=406 mm and thickness B=6.35 mm equal to at least about 185Mpa{square root}m.

Another important advantage of products according to the presentinvention is the fact that surprisingly, the value of K_(app(L-T)) asdetermined above is the same as or even higher when cold than its valueat ambient temperature. More precisely, this value evaluated at −54° C.is slightly greater than its value at ambient temperature. This isparticularly attractive, since −54° C. is approximately the typicaltemperature of structural members during the flight of a civil jetaircraft. It is known that the toughness decreases with the temperaturein some alloys in the 7xxx series. For example, it has been describedthat the toughness of plates made of 7475 T7651 reduces by 25%(determined from R curves on panels with thickness B=6 mm in the L-Tdirection) between about 20° C. and about −50° C. (see P. R. Abelkis etal., Proceedings of “Fatigue at Low Temperatures”, Louisville, Ky., May10, 1983, pages 257-273 (published by ASTM)). Under the same conditions,thick plates made of 7050 T7451 have a reduction in K_(IC) or K_(q) inthe L-T or T-L direction equal to at least 5% (see W. F. Brown et al.,Aerospace Materials Handbook, published by CINDAS (USAF CRDA HandbookOperation, Purdue University, 1997)). The inventors have also observed areduction in K_(IC) for thick plates made of 7075 T7351, 7475 T 7351,7475 T7651, and under-aged 7475; this reduction is of the order of 2% to10%. Thus, if it is known that the static mechanical characteristicsR_(P0.2) and R_(m) of alloys in the 7xxx series tend to increase whenthe temperature drops from about 20° C. to about −50° C., which improvessafety for the structure at this temperature, the drop in toughness ofalloys in the 7xxx series according to the state of the art has to betaken into account when sizing structural members. A product accordingto the instant invention does not have a significant reduction (in otherwords more than 2%) of the toughness at low temperature, and in somecases the toughness at low temperature is even slightly higher, i.e. upto about 2% higher or even up to about 5% higher.

As structural members for wing lower structures of aircrafts, productsaccording to the present invention advantageously replace structuralmembers made of alloys known as 2×24 alloys, for example a 2024 or 2324alloy. For example, rolled products according to the present inventionmay be thinner than about 10 mm and thus be used, for example, as afuselage skin. They may also be thicker than about 10 mm and thus beused as structural members such as for lower wing structures. Rolledproducts more than about 40 mm thick may be used, for example, for themanufacture of structural members by integral machining as describedbelow. Rolled products with a thickness of more than about 60 mm can beused, for example, for manufacturing stiffeners or frames, particularlyfor large capacity aircraft.

Products according to the present invention may be clad on at least oneface thereof if desired for any reason using methods and with alloysconventionally used to clad products made of Al—Zn—Cu—Mg type alloys.This is particularly attractive for plates used for manufacturingaircraft fuselage members that have to resist corrosion. One exemplarycladding alloy that can be used is 7072.

One particularly advantageous use of products according to the instantinvention is related to the concept of the integral structure inaeronautical construction. A large proportion of aircraft structures aresized as a function of a compromise between damage tolerance andresistance to static loads. Requirements for damage tolerance are, forexample specified in the article “Damage Tolerance Certification ofCommercial Aircraft” by T. Swift, ASM Handbook vol. 19 (1996), pp566-576. Sizing under static loads is explained for example in the book“Airframe Stress Analysis and Sizing” by M. Niu, Hong Kong CommilitPress Ltd, 1999, particularly pages 607 to 654. From the material pointof view, it is known that the damage tolerance of alloys in the 7xxxseries, and particularly their toughness, generally decreases when theiryield strength increases. This phenomenon leads to specialization ofalloys with high damage tolerance—particularly alloys in the 2xxxseries—for parts with very high tension stresses, knowing that thetolerance certification requires the acceptance of the presence ofcracks, and conversely alloys with a high yield strength—particularlyalloys in the 7xxx series, for parts with very high compressionstresses. In reality, parts with very high compression stresses such aswing upper surfaces and fuselage lower structures, are also subjected totension loads, which although they are not as high, make it necessaryfor the material to have a certain damage tolerance. Conversely, partssuch as wing undersurfaces and fuselage upper structures, in whichtension stresses are very high, require a certain minimum compressionstrength. Thus, it frequently happens that damage tolerance is acontrolling parameter for a part that is stressed essentially incompression, and vice versa. Thus for example, an increase in toughnessof x % for a constant yield strength as with the alloy according to thepresent invention may result in a corresponding weight saving, or evenbetter if the fact that allowing a high load on the part considered alsomakes it possible to reduce the weight of other components. Similarly,an increase in the yield strength equal to x %, for constant damagetolerance, can result in a weight saving of the order of x/3% to x %.For products according to the instant invention, x is typicallypreferably from about 15 to about 30%.

In an integral structure, continuity between the stiffeners and the skinmeans that damage tolerance becomes more critical than in a componentassembled by riveting. At a given stress, the stress intensity factorincreases strongly when a crack passes through the stiffener, since itmust be assumed that this stiffener will be necessarily cracked. Thepresent inventors have found that high toughness products, for a givenyield strength, are particularly suitable for manufacturing of integralstructures. In one particularly advantageous embodiment of this aspectof the present invention, fuselage lower structure panels and wing skinsare made by integral machining of products according to one of theprevious embodiments. Such products, and particularly thin plates, to bemachined are advantageously at least 40 mm thick; this value alsodepends on the type of aircraft and particularly its size. According tothe observations made by the inventors, a weight saving of the sameorder of magnitude as the improvement in toughness, namely about 10%,can be achieved compared with an integral structure made from a typeAA7475 alloy according to the state of the art. More precisely, aproduct according to the invention, with a yield strength R_(P0.2(L)) atmid-thickness equal to at least about 540 MPa and a toughnessK_(app(LT)) measured on an M(T) type specimen with a width W of 16inches (about 406 mm) equal to at least about 140 Mpa{square root}m, canbe used to make structural members for aircraft such as a wing skinmember with a weight saving equal to at least 10% compared with the samepart with the same shape and size made from a 7475 alloy according tothe state of the art and typically having an R_(P0.2(L)) atmid-thickness equal to 475 MPa, and an K_(app(LT)) measured on an M(T)type specimen with a width W of 16 inches (about 406 mm) equal to 125Mpa{square root}m.

The inventors have observed that refining the grain to a lower levelthan is accepted in normal practice during casting can give aparticularly attractive compromise between properties, particularly fortoughness. The use of a refining agent made of TiC (for example additionof an A13% Ti0.15% C wire) in controlled doses is particularlybeneficial, the solidification germ obtained with this approach having adifferent compromise between germination and growth than is possiblewith germs obtained for example by refining with A15% Ti1% B (in otherwords a TiB₂ type germ). The level of this refining may be quantified bythe quantity of C added, since it indirectly corresponds to the quantityof added solidification germs and the quantity of free Ti (not combinedwith C) into the alloy. Although the stoechiometry of the germ is notdefinitively quantified, it can be considered that the germ comprisesTiC, each C atom combining with a Ti atom to form the said germs.

There are different types of refining agents Al-x % Ti-y % C, in generalexcess Ti being added compared with C. The quantity of added germs ispreferably proportional to the quantity of refining agent (in kilograms)added per ton of liquid metal multiplied by y/o, in other wordsproportional to A (number of kilograms of refining agent added per tonof metal)×y %.

Thus, for example, for the addition of 2 kg/t of Al-3% Ti-0.15% C, theaddition of germs can be quantified by specifying 3 g/t of added C(2×0.0015 kg/t).

There are other means of adding Al-3% Ti-0.15% C to arrive at theaddition of the same quantity of germs, for example by adding twice asmuch refining agent with half the concentration of C.

In one advantageous embodiment of this invention, a refining agentcomprising titanium and carbon is also added such that the added carbonquantity is preferably between 0.4 and 3 g/t of carbon, more preferablybetween 0.6 and 2 g/t and such that the total content of Ti in the finalproduct is between 50 and 500 ppm (by weight) and preferably between 150and 300 ppm.

Other advantageous embodiments are described by the claims.

In the following example, advantageous embodiments of the invention aredescribed for illustration purposes. These examples are in no waylimitative.

EXAMPLE 1

An alloy N was made for which the chemical composition complies with theinvention. The liquid metal was treated firstly in a holding furnace byinjecting gas using an IRMA® type of rotor, and then in an Alpur® typeof ladle, these two trademarks belonging to the inventors. Refining wasdone in line, in other words in the channel between the holding furnaceand the Alpur® ladle, with 1.1 kg/tonne of Al-3% Ti-0.15% C wire (9.5 mmdiameter). An industrial sized rolling ingot was cast. It was relaxedfor 10 h at 350° C.

The product thus cast was homogenised after scalping for 15 hours at atemperature between 471° C. and 482° C. (between 880° F. and 900° F.)and then hot rolled to a thickness of 5 mm (0.2 inches). The rollingstart temperature was 450° C. (840° F.) and the rolling end temperaturewas 349° C. (660° F.). Plates with width 178 mm (7 inches) and length508 mm (20 inches) were sampled. These coupons were solution heattreated in a salt bath furnace for 1 hour at 472° C. and then quenchedin water and tensioned to obtain a permanent deformation of 2%. Thecoupons thus obtained were then subjected to a two-step artificial agingtreatment, the first step being 6 hours at 105° C., the second stepbeing 18 hours at 155° C., in order to reach the peak of mechanicalproperties.

Using a similar process, sheet with a thickness of 6 mm and 3.2 mm inalloy Y was elaborated.

Plates made of 2xxx type alloys (references E and F outside the scope ofthe invention) were also produced according to the following process:

The alloy was cast firstly by treating the liquid metal in a holdingfurnace by injecting gas using an IRMA® type of rotor, and then in anAlpur® type of ladle. Refining was done in line, in other words in thechannel between the holding furnace and the Alpur® ladle, with 0.7kg/tonne of AT5B wire (9.5 mm diameter). The cast rolling ingots werestress relieved for 10 hours at 350° C. These rolling ingots were thenhomogenised for 12 hours at 500° C., then hot rolled (end of rollingtemperature between 230 and 255° C.) to a thickness of 6 mm. A solutionheat treatment was then carried out in a salt bath furnace for 1 hour at500° C. on the 600 mm by 200 mm coupons. This operation was followed byquenching in cold water at about 20° C. and stretching with a permanentset of 2% (temper T351).

Rolling ingots made of a 7xxx alloy according to prior art were alsocast (reference G), in the same founding device as plates made with 2xxxalloy described above. The resulting rolling ingot was homogenised for24 hours at 470° C. and then 24 hours at 495° C., then hot rolled (endof rolling temperature between 230 and 255° C.) to a thickness of 6 mm.A solution heat treatment of 1 hour was then carried out at 450° C. in asalt bath furnace on a 600 mm by 200 mm coupon. This operation wasfollowed by quenching in water and stretched with a permanent set of 2%.The coupon was then subjected to artificial aging treatment for 5 hoursat 100° C. then 6 hours at 155° C., in order to achieve the peakmechanical properties (temper T6).

A rolling ingot made of an AA7475 alloy was also cast (reference H)according to conventional processes according to prior art. The rollingingot thus obtained was homogenised for 9 hours at 480° C., and then hotco-rolled at a temperature of about 270° C. with a 7072 cladding plate,until a sheet with a thickness of 4.5 mm was obtained. The 7072 claddingaccounts for about 2% of the final thickness. The product thus obtainedwas solution heat treated in a salt bath furnace for 45 minutes at 478°C., then quenched in water at a temperature of about 20° C., and thenstretched with a permanent set of 2%. It was then subjected to atwo-step artificial aging operation for 4 hours at 120° C., then 24hours at 162° C. (temper T76).

The chemical compositions of the N, Y, E, F, G and H alloys measured ona spectrometry slug taken from the casting runner, are given in table 1:TABLE 1 Chemical composition Alloy Si Fe Cu Mn Mg Zn Zr Cr N (invention)0.05 0.06 2.05 — 1.64 7.08  0.08 — Y (invention) 0.04 0.05 2.16 — 1.806.76  0.09 — E (2024A) <0.06 0.06 4.12 0.4 1.37 — — — G 0.05 0.08 1.47 —1.56 4.27  0.11 — H (7475) 0.03 0.06 1.5 — 2.22 5.73 —  0.21 Cladding0.15 0.35 <0.02 <0.05 <0.10 1.05 <0.03 <0.03 (7072)

The ultimate tensile strength R_(m) (in MPa), the tensile yield strengthat 0.2% elongation R_(P0.2) (in MPa) and the elongation at failure A (in%) were measured using a tensile test according to EN 10002-1.

The results of measurements of the static mechanical characteristics fortemper T6 for plates N and Y according to the invention, and for temperT351 for plates E, F and G according to prior art, are shown in table 2:TABLE 2 Static mechanical characteristics Thick- L direction TLdirection ness R_(p0.2) R_(p0.2) Plate [mm] R_(m)[MPa] [MPa] A[%]R_(m)[MPa] [MPa] A[%] N 5.08 539 508 13.9 541 495 13.9 Y 6 557 530 13.9555 519 13.6 E 6.35 482 365 22.8 466 319 23.5 G 6.35 435 373 15.1 436366 14.8 H 4.6 475 414 13.3 484 414 12.5

It can be seen that the ultimate tensile strength and tensile yieldstrength of the plate according to the invention in both measureddirections is very much higher than the corresponding values for platesmade of a 2xxx alloy. The elongation of the plate according to theinvention is lower than that of plate E, but is sufficient for thetarget applications. Compared with 7xxx alloys according to prior art Gand H, the alloy according to the invention has a significantly improvedultimate strength and yield strength for a comparable elongation.

Plates N, E, F, G and H were evaluated to determine the toughnessmeasured by determination of the stress intensity factors K_(e0) orK_(app) according to standard ASTM 561; this determination was made inthe T-L direction, on C(T) test pieces with W=127 mm (5 inches) andB=5.5 mm.

The results are shown in Table 3 below. TABLE 3 Measurements of K_(app)Plate K_(app) [MPa{square root}m] N (invention) 107 E (2024 A) 105 G 97H (7475 cladded) 87

Sheet Y in thickness 6 mm had a fracture toughness K_(app) of 150MPa{square root}m (for W=760 mm) or 134 Mpa{square root}m (for W=406 mm)in the L-T direction, and of 128 Mpa{square root}m (for W=760 mm) or 110Mpa{square root}m (for W=406 mm) in the T-L direction.

The value of K_(app) for the plate according to the invention is muchgreater than the value for plates made of 7xxx alloy according to priorart, and is of the same order of magnitude as for plates made of 2xxxalloy.

The fatigue behaviour was also tested according to ASTM standard E 647,measuring the crack propagation rate in plate N in comparison to platesE, F and G. The test pieces used were of the C(T) type, where W is 76.2mm (3 inches).

The crack propagation rate results dA/dN for ΔK equal to 10 Mpa{squareroot}m, then 30 Mpa{square root}m were measured; the value of ΔK for apropagation rate of 100 μinch/cycle (or 0.24 μm/cycle) was measured. Thecomparative results are given in table 4. Sheet Y had a thickness of 6mm. TABLE 4 Fatigue results da/dN (10) da/dN (30) ΔK T-L T-L at 100μinch/cycle Plate (10⁻⁴ mm/cycles) (10⁻⁴ mm/cycles) [MPa{square root}m]N (invention) 1.4 29 27.5 Y (invention) 1.3 33 27 E (2024 A) 1.4 30 27 G1.1 38 25.9

The plate according to the invention behaves just is well in fatigue asplates according to prior art.

Another sheet Y with a thickness of 3.2 mm had the following properties:da/dN(10)(T-L)=1.710⁻⁴ mm/cycles, da/dN(30)(T-L)=30 10⁻⁴ mm/cycles, AKat 100 μinch/cycle=28.3 Mpa{square root}m.

EXAMPLE 2

An alloy M with chemical composition complying with the invention wasproduced.

For comparison, a plate made of 2324 alloy according to prior art(reference 1) was produced according to a conventional casting process.

The chemical compositions of alloys M and I measured on a spectrometryslug taken from the casting runner, are given in the following table:TABLE 5 Chemical composition Alloy Si Fe Cu Mn Mg Zn Zr M 0.05 0.06 2.05— 1.64 7.08 0.08 I (AA2324) <0.10 <0.12 3.8-4.4 0.3-0.9 1.2-1.8 <0.20<0.05

After scalping, rolling ingots made of alloy M were homogenised for 15hours at 479° C., and then slowly cooled to 420-440° C. and rolled to athickness of 25.4 mm. The outlet temperature from the hot rolling millwas 354° C., which is significantly lower than the value normally usedfor this type of product.

The plates thus obtained were then subjected to a solution heattreatment at 479° C. for 4 hours (total time, about ⅓ of which is spentin the temperature increase), and were then quenched and tensioned suchthat the resulting permanent deformation is 2%. The plates were thensubjected to an artificial aging treatment for 8 hours at 160° C.

The ultimate tensile strength R_(m) (in MPa) the tensile yield strengthat 0.2% elongation Rp_(0.2) (in MPa) and the elongation at fracture A(in %) were measured using a tensile test according to EN 10002-1 forthe plate according to the invention, and for the plate according toprior art. The corresponding results are shown in Table 6 below.

Alloy I (AA2324) was subjected to a conventional procedure to obtain aplate made of AA 2324 alloy, 25.4 mm thick in the T39 temper, in otherwords a homogenisation step followed by a hot rolling step, thensolution heat treatment and quenching, followed by cold working of about9%, and controlled stretch with a permanent set of between 1.5 and 3%.TABLE 6 Static mechanical properties L direction Thickness RmR_(P0.2(L)) Plate [mm] [MPa] [MPa] A [%] M 25.4 570 540 12.3 I 25.4 490470 14

It was observed that the ultimate strength and the yield strength of theplate according to the invention are significantly higher than thecorresponding values of plate I usually used for these applications, andfor quite comparable elongations.

The toughness was also evaluated for plates M and I, measured bydetermining the critical stress intensity factors K_(C) and K_(CO) orK_(app), according to the ASTM standard 561; this determination was madeat ambient temperature in the L-T direction on M(T) test pieces withB=6.35 mm (0.25 inches) and W=406.4 mm (16 inches), and on C(T) testpieces with B=7.6 mm (0.3 inches) and W=127 mm (5 inches). K_(app) wasalso determined on a C(T) test piece with B=7.6 mm and W=127 mm in theL-T direction at a temperature of −54° C. The results are given in table7 below. TABLE 7 Measurements of K_(C) and K_(app) K_(C(L-T)) testK_(app(L-T)) test K_(app(L-T)) test K_(app(L-T)) K_(app(L-T)) piece M(T)piece M(T) piece C(T) test piece test piece [MPa {square root}m] [MPa{square root}m] [MPa {square root}m] C(T) M(T) ambient ambient ambient[MPa {square root}m] [MPa {square root}m] Plate temperature temperaturetemperature at −54° C. at −54° C. M 199 140 118 124 126 I 177 121 96 99—

It was found that the alloy according to the invention had bettertoughness than the conventional alloy I under all conditions. And alsosurprisingly, the alloy according to the invention had a value ofK_(app(LT)) that was of the same order at −54° C. as it is at ambienttemperature.

Plates M and I were also tested for fatigue strength along the Ldirection, using the following two protocols taken from ASTM standard E466:

-   -   1) A notched test piece 5 mm thick, 38.1 mm wide and 254 mm long        was used with two circular notches with radius 43.2 mm machined        symmetrically about the centre of the test piece at a distance        of 12.7 mm from the centre. The test is made according to ASTM        standard E 466, by applying a cyclic stress such that the        maximum stress is equal to 270 MPa and the minimum stress is        equal to 27 MPa (R=0.1), at a frequency of 15 Hz.    -   2) A “double hole” test piece 2.54 mm thick, 25.4 mm wide and        209 mm long, was used, with two circular holes with diameter 4.8        mm located on the median line of the test piece, at equal        distance from the centre of the test piece, and with centres at        a spacing of 19 mm. The test is made according to ASTM standard        E 466, by applying a cyclic stress such that the maximum stress        is equal to 140 MPa, and the minimum stress is equal to 14 MPa        (R=0.1), at a frequency of 15 Hz.

This test was carried out on five test pieces for each protocol, and thelogarithmic average of 5 tests was calculated.

The results of these two test protocols on two plates M and I are givenin Table 8 below: TABLE 8 Fatigue results Number of cycles (Log Numberof cycles (Log average on 5 tests) average on 5 tests) Plate “Notched”test piece “Double hole” test piece M 299 213 330 737 I 181 402 337 730

The variability of this test is generally fairly high, but it isobserved on this test that the plate according to the invention and theplate usually used have the same orders of magnitude in terms of fatiguelife.

The fatigue behaviour was also tested according to ASTM standard E 647,by measuring the crack propagation rate in plate M in comparison withplate I. The test pieces used were of the C(T) type where B is equal to9.52 mm (0.375 inches) and W is equal to 101.6 mm (4 inches).

The crack propagation rate curve was plotted as a function of ΔK, andthe value of ΔK at a rate of 2.54 μm/cycle (10⁻⁴ inch/cycle) wasmeasured; the comparative results are given in table 9. TABLE 9 Fatigueresults Values of ΔK for a given crack rate ΔK at 2.54 μm/cycle Plate(10⁻⁴ inch/cycle) [MPa{square root}m] M 30.8 I 26.8

Finally, the exfoliation corrosion behaviour of plates in this test wasevaluated according to ASTM standard G34; this test was done on thesurface and at mid-thickness under conditions adapted to 7xxx alloys forplate M according to the invention, and under conditions adapted to 2xxxalloys for plate I. Sample M according to the invention was classifiedEA, both at the surface and at mid-thickness, while sample I accordingto prior art was classified EA at the surface and EB at mid-thickness.Therefore, the performance of the plate according to the invention interms of exfoliation corrosion is at least as good, if not better, thanthe plate according to prior art.

It is observed that plate M is better for static mechanicalcharacteristics, K_(app), fatigue resistance and for the crackpropagation rates.

EXAMPLE 3

An alloy P similar to alloy M in example 2 was produced. A manufacturingprocedure similar to that for example 2 was used to make thickintegrally hot rolled plates from this alloy (input temperature 420-440°C.), with a thickness of 75 mm.

After solution heat treatment and quenching as indicated in example 2,the plates were subjected to annealing processes with the following twosteps:

-   -   First step: temperature increase at 30° C./hour up to 120° C.        and hold for 6 hours at this temperature of 120° C.    -   Second step: temperature increase at 15° C./hour up to 160° C.        and hold for 5 hours (process A), 10 hours (process B) or 15        hours (process C) at this temperature of 160° C.

The values of K_(app(LT)) were determined on type C(T) test pieces withW=127 mm and B=7.6 mm.

Table 10 summarises the main mechanical characteristics obtained: TABLE10 R_(P0.2(L)) R_(m(L)) A_((L)) K_(IC(L-T)) K_(app(L-T)) Process [MPa][MPa] [%] [MPa{square root}m] [MPa{square root}m] A 542 561 9.7 30.157.1 B 525 549 10.2 32.8 63.2 C 507 537 11.3 34.6 72.5

Additional advantages, features and modifications will readily occur tothose skilled in the art. Therefore, the invention in its broaderaspects is not limited to the specific details, and representativedevices, shown and described herein. Accordingly, various modificationsmay be made without departing from the spirit or scope of the generalinventive concept as defined by the appended claims and theirequivalents.

Units listed herein and in the following claims can be expanded to coverclose values so long as one or more inventive concepts described hereinare maintained.

All documents referred to herein are specifically incorporated herein byreference in their entireties.

As used herein and in the following claims, articles such as “the”, “a”and “an” can connote the singular or plural.

1. A work-hardened product comprising an alloy having a composition (%by weight)of: Zn 6.7-7.3% Cu 1.9-2.5% Mg 1.0-2.0% Zr 0.04-0.15% Fe ≦0.15Si ≦0.15 other elements ≦0.05 each and ≦0.15 total, remainder aluminium,wherein Mg/Cu<1, said product being treated by solution heat treatment,quenching, cold working and artificial aging.
 2. A product according toclaim 1, wherein Zr≦0.12%.
 3. A product according to claim 1, wherein Zrcontent is from about 0.07 to about 0.09%.
 4. A product according toclaim 1, wherein the Cu content is from about 2.0 to about 2.3%.
 5. Aproduct according to claim 1, wherein the Mg content is from about 1.5to about 1.8%.
 6. A product according to claim 1, wherein Mg/Cu≦0.80. 7.A product according to claim 1 having a thickness of at most about 20mm, wherein (a) K_(app(L-T)) (measured at ambient temperature on a C(T)type test piece with W=127 mm and B=5.5 mm)>100 MPa{square root}m (b)R_(p0.2(L))>500 MPa.
 8. A product according to claim 1, whereinR_(m(L))>500 MPa.
 9. A product according to claim 1, whereinR_(P0.2(L))>about 460 Mpa.
 10. A plate with a thickness of at most about20 mm comprising a product according to claim 2, having at least oneselected from the group consisting of: (a) R_(m(L))>520 MPa andR_(P0.2(LT))>515 MPa; (b) R_(P0.2(L))>500 MPa and R_(P0.2(LT))>480 MPa;and (c) R_(P0.2(L))>500 MPa and K_(app(TL))>100 Mpa{square root}m(measured at ambient temperature on a C(T) type test piece with W=127 mmand B=5.5 mm).
 11. A plate according to claim 10, wherein da/dn in theT-L direction is at least about 28×10−4 mm/cycles, determined at ΔK=30Mpa{square root}m.
 12. A sheet or plate with a thickness from about 1 toabout 10 mm comprising a product according to claim 1, whereinK_(C(T-L)), measured on a test piece with W=760 mm, is at least about130 Mpa{square root}m, and/or wherein K_(C(L-T)), measured on a testpiece with W=760 mm, is at least about 160 Mpa{square root}m.
 13. Aplate according to claim 10, wherein said plate has been hot rolledonly.
 14. A plate with thickness of at least about 10 mm comprising aproduct according to claim 2, wherein said plate has been heat treatedfor at least 4 hours at a temperature of from about 470 to about 480° C.15. A plate with thickness of at most about 10 mm comprising a productaccording to claim 2 wherein said plate has been solution heat treatedfor at least 1 hour at a temperature of from about 470 to about 480° C.16. A plate with a thickness of at least about 20 mm comprising aproduct according to claim 2 having at least two characteristicsselected from the group consisting of: (a) R_(m(L))>540 MPa; (b)R_(P0.2(L))>535 MPa; (c) K_(app(LT))>100 Mpa{square root}m (measured atambient temperature on a C(T) type test piece with W=127 mm and B=7.6mm); (d) ΔK at a crack propagation rate of 2.54 μm/cycle>28 Mpa{squareroot}m; and (e) K_(IC)(L-T)>28 MPa{square root}m.
 17. A plate with athickness of at least 20 mm comprising a product according to claim 2,having at least two characteristics selected from the group consistingof: (a) yield strength R_(p0.2(L)) equal to at least 520 MPa; (b)K_(app(L-T)) (measured at ambient temperature according to ASTM 561 on aC(T) type test piece with W=406 mm and thickness B=6.35 mm) equal to atleast 130 Mpa{square root}m; and (c) K_(C(L-T)) (measured at ambienttemperature on a C(T) type test piece with W=406 mm and thickness B=6.35mm) equal to at least 185 Mpa{square root}m.
 18. A plate comprising aproduct of claim 1 that is clad on at least one face thereof.
 19. Aproduct according to claim 1 wherein said product has been produced froma liquid metal to which a refining agent containing titanium and carbonhas been added, such that the carbon quantity added to said liquid metalis from about 0.4 to about 3 g/t of carbon, and such that the totalcontent of Ti in the product is from about 50 to about 500 ppm (byweight).
 20. A structural member suitable for aircraft comprising atleast one product according to claim
 1. 21. An integral structuresuitable for aircraft comprising at least one product according toclaim
 1. 22. A fuselage skin comprising a rolled product comprising aproduct according to claim 2 and having a thickness of at most about 10mm.
 23. A method for machining aircraft structural members comprisingusing a product according to claim 1 having a thickness of at leastabout 40 mm.
 24. A method for making aircraft stiffeners or framescomprising using a product according to claim 1 with a thickness of atleast about 60 mm.
 25. A product wherein the value of K_(app(L-T))measured according to ASTM 561 on a C(T) type test piece with W=406 mmand thickness B=6.35 mm when measured at −54° C. is not more than about2% less than said value at ambient temperature.
 26. A product of claim25, wherein said value at −54 C is the same or slightly greater thansaid value at ambient temperature.
 27. A product with a yield strengthR_(P0.2(L)) at mid-thickness equal to at least about 540 MPa and atoughness K_(app(LT)) measured on an M(T) type specimen with a width Wof 16 inches (about 406 mm) equal to at least about 140 Mpa{squareroot}m, said product comprising a structural member for aircraft andhaving a weight at least 10% less than a structural member with the sameyield strength and toughness comprising a 7475 alloy.
 28. A product ofclaim 27 wherein said product comprises an alloy having a composition (%by weight) of: Zn 6.7-7.3% Cu 1.9-2.5% Mg 1.0-2.0% Zr 0.04-0.15% Fe≦0.15 Si ≦0.15 other elements ≦0.05 each and ≦0.15 total, remainderaluminium, wherein Mg/Cu<1.
 29. An integral structure comprising a plateaccording to claim
 16. 30. An integral structure of claim 29 that is atleast about 40 mm in thickness.
 31. An integral structure comprising aplate according to claim
 17. 32. An integral structure according toclaim 31 that is at least about 40 mm in thickness.
 33. A structuralmember comprising a product of claim
 2. 34. A structural member of claim33 that is at least about 10 mm in thickness.
 35. A structural membercomprising a product of claim
 3. 36. A structural member of claim 35that is at least about 10 mm in thickness.
 37. A structural membercomprising a plate according to claim 14.